Systems and methods for rotor/wing aircraft

ABSTRACT

Systems and methods for rotor/wing aircraft are disclosed. In one embodiment, an aircraft includes an airframe, a high-lift canard and tail, a rotor/wing, a propulsion system, and a drive assembly. The drive assembly, which may include a radial inflow turbine, is configured to extract work from the propulsion system to selectively rotate the rotor/wing assembly thus enabling the aircraft to conduct rotary-wing flight, fixed wing flight as well as smoothly transition between the two modes of flight.

CROSS REFERENCE TO RELATED APPLICATIONS

This patent application is a continuation-in-part application ofcommonly-owned U.S. patent application Ser. No. 11/265,655 entitled“Rotor-Wing Aircraft Having an Adjustable Tail Nozzle” filed on Nov. 2,2005, now U.S. Pat. No. 7,395,988, which application and issued patentis incorporated herein by reference.

FIELD OF THE INVENTION

This disclosure relates to systems and methods for rotor/wing aircraft.

BACKGROUND

Higher performance rotary wing aircraft are sought. A stopped-rotor (orrotor/wing) aircraft like that shown in U.S. Pat. No. 3,327,969 offersthe ability to hover like a helicopter plus the promise of achieving thehigh speeds of a fixed-wing aircraft by stopping the rotor's rotationwhile in flight and allowing it to act as a fixed wing.

Another prior art example of a stopped-rotor aircraft, the CanardRotor/Wing (CRW) concept (U.S. Pat. No. 5,454,530), combines areaction-driven, stoppable-rotor with a high-lift canard and tail.Exhaust gas from a common power plant (i.e. gas turbine engine) providesdirect thrust required for fixed-wing flight and is routed to the bladetips to power the reaction-driven rotor for rotary-wing flight. Togetherthe canard and tail provide all of the aircraft's lift during transitionbetween rotary-wing and fixed-wing flight, thereby allowing the rotor tobe unloaded during starting and stopping.

It is known that all reaction driven rotors have their own uniqueinefficiencies. Coriolis losses associated with accelerating thepropulsive gases radially as the rotor blade spins consumes up to 40% ofpower available. Thus, stopped-rotor aircraft designs are sought thatare more efficient than the reaction-driven, rotor/wing system of theCRW.

SUMMARY OF THE INVENTION

Embodiments of systems and methods for rotor/wing aircraft aredisclosed.

In one embodiment, an aircraft includes an airframe, a rotor/wing, ahigh-lift canard and tail, a propulsion system, and a power-takeoffassembly that enable the aircraft to conduct rotary-wing flight, fixedwing flight and smoothly transition between the two modes. In someimplementations, the aircraft may further include a rotatable aft nozzlethat is configured to controllably vector engine thrust so as tosupplement the lift required from the rotor/wing during rotary-wingflight.

In another embodiment, a drive assembly is replaced by a radial inflowturbine configured to extract work from the propulsion system for thepurpose of rotating the rotor/wing to produce lift during rotary-wingflight. The radial inflow turbine can be selectively by-passed for thepurpose of producing thrust during fixed-wing flight. A high-lift canardand tail enable smooth transition between the two modes. In furtherembodiments, a rotatable aft nozzle is included in the assembly with theradial inflow turbine.

The features, functions, and advantages that will be discussed can beachieved independently in various embodiments of the present disclosureor may be combined in yet other embodiments further details of which canbe seen with reference to the following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the present disclosure are described in detail below withreference to the following drawings.

FIG. 1 is a side sectional view of an aircraft in a fixed-wing mode ofoperation in accordance with another embodiment of the presentdisclosure.

FIGS. 2 and 3 are side sectional views of the aircraft of FIG. 1 inconversion and rotary-wing flight modes of operation, respectively.

FIG. 4 is a side sectional view of an aircraft in a rotary-wing mode ofoperation in accordance with yet another embodiment of the presentdisclosure.

FIG. 5 is a side sectional view of an aircraft in a fixed-wing mode ofoperation in accordance with another alternate embodiment of the presentdisclosure.

FIGS. 6 and 7 are side sectional views of the aircraft of FIG. 5 inconversion and rotary-wing flight modes of operation, respectively.

FIG. 8 is a perspective of an aircraft according to an embodiment of thepresent disclosure.

FIG. 9 is a side view of the aircraft of FIG. 8.

FIG. 10 is a perspective of a radial inflow turbine according to anotherembodiment of the present disclosure.

FIG. 11 is a perspective of the radial inflow turbine of FIG. 10 shownwithout half of a body of the radial inflow turbine.

FIG. 12 is a cross section taken along lines 5-5 of FIG. 10.

DETAILED DESCRIPTION

Many specific details of certain embodiments of the disclosure are setforth in the following description and in FIGS. 1-12 to provide athorough understanding of such embodiments. One skilled in the art,however, will understand that the present disclosure may have additionalembodiments, or that the present disclosure may be practiced withoutseveral of the details described in the following description.

FIG. 1 is a side view of an aircraft 100 in a fixed-wing mode 150 ofoperation in accordance with an embodiment of the present disclosure. Inthis embodiment, the aircraft 100 includes a fuselage 102 of an airframedefining a cockpit portion 104. A canard 106 extends outwardly from eachlateral side of the fuselage 102 (one visible). The aircraft 100 furtherincludes a tail assembly 110 having a horizontal tail surface 112 thatextends outwardly from each lateral side of the fuselage 102 (onevisible), and a vertical tail surface 114 that extends upwardly fromeach horizontal tail surface 112 (one visible).

As further shown in FIG. 1, the aircraft 100 includes a propulsionsystem 120 having at least one engine 122 disposed within an aft portionof the fuselage 102 proximate the tail assembly 110. An engine exhaustflow from the engine 122 is directed through a nozzle assembly 126 toprovide a primary thrust 124 for propelling the aircraft 100 in thefixed-wing mode 150 of operation. In the fixed-wing mode 150 ofoperation, a canard lift 107 is generated by each canard 106, and a taillift 113 is generated by each horizontal tail surface 112.

A primary shaft 128 extends generally forwardly within the fuselage 102from the engine 122 to a gearbox assembly 130. The primary shaft 128 maybe driven by the engine 122 using a conventional power-takeoffconfiguration generally known in the art. A secondary shaft 132 extendsgenerally upwardly from the gearbox assembly 130 to a rotor/wingassembly 134. In the fixed-wing mode 150 of operation shown in FIG. 1,the rotor/wing assembly 134 is held in a fixed position suitable forforward flight.

FIG. 2 shows the aircraft 100 of FIG. 1 in a conversion mode 152 ofoperation. In this embodiment, the primary wings (or canard) 106 and thetail assembly 110 are in the process of transitioning from thefixed-wing mode 150 of operation to a rotary-wing mode 154 (describedbelow with respect to FIG. 3). More specifically, in some embodiments,each canard 106 and tail assembly 110 is pivotable between a firstposition (FIG. 1) configured to provide a lift force in the fixed-wingmode 150 of operation, and a second position (FIG. 3) configured tominimize a downward force in the rotary-wing mode 154 of operation.

In addition, rotation of the secondary shaft 132 has initiated rotationof the rotor/wing assembly 134 in a rotational direction 136. Comparisonof FIG. 2 with FIG. 1 shows that the primary thrust 124 from the engine122 is reduced in comparison with the fixed-wing mode 150 of operation.In the conversion mode 152 of operation, the lift from the canards 106and the tail assembly 110 may continue to provide 100% of the liftrequired to maintain flight to allow the rotor/wing assembly 134 tostart and stop while unloaded. The vertical tail surfaces 114 of thetail assembly 110 provide antitorque forces during the conversion mode152 of operation.

The aircraft 100 is shown operating in the rotary-wing mode 154 ofoperation in FIG. 3. In this embodiment, the canards 106 and the tailassembly 110 are fully pivoted from their respective positions in thefixed-wing mode 150 of operation (e.g. approximately 90 degrees) shownin FIG. 1. In addition, the rotational velocity of the primary shaft 128has increased such that a primary lift component 138 from the rotor/wingassembly 134 is sufficient to support the aircraft 100 in verticalflight (ascent or descent) or hover as desired. Further, the primarythrust 124 from the engine 122 may be approximately zero, and a portionof the engine exhaust flow may be used as a yaw control (or anti-torque)flow 125 to counteract a torque generated by the rotation of therotor/wing assembly 134. For example, the yaw control flow 125 may beexhausted through a yaw control nozzle to provide a control force havinga laterally-directed component that counteracts the torque generated bythe rotation of the rotor/wing assembly 134. More specifically, in someembodiments, the yaw control flow 125 may be directed through a yawcontrol system such as the NOTAR® system, or any other suitable yawcontrol system.

FIG. 4 is a side view of an aircraft 160 in a rotary-wing mode 162 ofoperation and is shown as an enhancement to the first embodiment of thepresent disclosure. Many of the components of the aircraft 160 shown inFIG. 4 are similar or identical to the components of the aircraft 100described above with respect to FIGS. 1 through 3, and therefore, forthe sake of brevity, a description of such components will not berepeated. In this embodiment, however, the aircraft 160 includes avectorable nozzle 166 that is configured to controllably vector anengine thrust component 164 from approximately horizontal in thefixed-wing mode 150 to approximately vertical in the rotary-wing mode154 (FIG. 4). Thus, in addition to the primary lift components 138generated by the rotor/wing assembly 134, the aircraft 160 may also relyupon the engine thrust component 164 during the rotary-wing mode 162 ofoperation. When the aircraft 160 returns to the fixed-wing mode 150 ofoperation (FIG. 1), the vectorable nozzle 166 may be returned to agenerally aftward-facing direction to provide the generally horizontalthrust 124 needed for forward flight as shown in FIG. 1.

FIGS. 5 through 7 are side sectional views of an aircraft 200 inaccordance with another embodiment of the present disclosure. Morespecifically, in FIG. 5, the aircraft 200 is shown in a fixed-wing mode250 of operation. The aircraft 200 includes a fuselage (or airframe) 202defining a cockpit portion 204, and having canards 206 that extendoutwardly from each lateral side of the fuselage 202 (one visible). Theaircraft 200 further includes a tail assembly 210 having a horizontaltail surface 212 that extends outwardly from each lateral side of thefuselage 202 (one visible), and a vertical tail surface 214 that extendsupwardly from each horizontal tail surface 212 (one visible).

As further shown in FIG. 5, the aircraft 200 includes a propulsionsystem 220 having at least one engine 222 coupled to (or disposedwithin) a dorsal (or upper) portion of the fuselage 202. An engineexhaust flow from the engine 222 is directed through a nozzle assembly226 to provide a primary thrust 224 for propelling the aircraft 200 inthe fixed-wing mode 250 of operation. In the fixed-wing mode 250, acanard lift 207 is generated by each canard 206, and a tail lift 213 isgenerated by each horizontal tail surface 212.

A radial inflow turbine 240 is positioned aft of the engine 222. In someembodiments, the radial inflow turbine 240 is of a type as generallydescribed more fully below with respect to FIGS. 10 through 12. Morespecifically, in particular embodiments, the aircraft 200 includes twoengines 222 in a laterally-spaced (or side-by-side) arrangement (onevisible), and the radial inflow turbine 240 is configured as shown inFIGS. 10 through 12. A drive shaft 242 extends generally upwardly fromthe radial inflow turbine assembly 240 to a rotor/wing assembly 244. Inthe fixed-wing mode 250 of operation shown in FIG. 5, the rotor/wingassembly 244 is held in a fixed position suitable for forward flight.

FIG. 6 shows the aircraft 200 of FIG. 5 in a conversion mode 252 ofoperation. In this embodiment, the canards 206 and the tail assembly 210have started pivoting as described above with respect to the embodimentshown in FIG. 2. Rotation of the drive shaft 242 has initiated rotationof the rotor/wing assembly 244 in a rotational direction 246. In theconversion mode 252 of operation, the lift from the canards 206 and thetail assembly 210 may continue to provide 100% of the lift required tomaintain flight during transition to allow the rotor/wing assembly 244to start and stop while unloaded. The vertical tail surfaces 214 of thetail assembly 210 provide antitorque forces during the conversion mode252 of operation.

FIG. 7 shows the aircraft 200 in the rotary-wing mode 254 of operation,in which the canards 206 and the tail assembly 210 have fully pivoted asdescribed above with respect to the embodiment shown in FIG. 3. In thisembodiment, the rotational velocity of the drive shaft 242 has increasedsuch that the rotor/wing assembly 244 is rotating at sufficient velocityto generate a primary lift component 238 from the rotor/wing assembly244 sufficient to support the aircraft 200 in vertical flight (ascent ordescent) or hover. A portion of the engine exhaust flow may be used as ayaw control (or anti-torque) flow 225 to counteract a torque generatedby the rotation of the rotor/wing assembly 244 using a suitable yawcontrol system.

Referring to FIG. 8, an aircraft according to another embodiment of thepresent disclosure is designated in its entirety by reference number 10.In this embodiment, the aircraft 10 has an airframe, generallydesignated by 12, which includes a fuselage 14 having a nose or forwardend 16 and a tail or aft end 18. The aircraft 10 further includes atleast two primary fixed wings or canards 20 extending laterally from thefuselage. Each primary fixed wing 20 has a wing tip 22 opposite thefuselage 14. The aircraft 10 may also include a rear set of fixed wings24. Each rear fixed wing 24 has a wing tip 26 opposite the fuselage 14.

The specific dimensions of the components of the aircraft 10 may varyconsiderably from embodiment to embodiment. For example, although thefuselage 14 may have other lengths extending between the forward end 16and the aft end 18 without departing from the scope of the presentdisclosure, in some embodiments the fuselage 14 has a length of betweenabout 60 feet and about 70 feet. Similarly, in some embodiments theaircraft 10 may have a canard span of between about 35 feet and about 45feet, and in further embodiments, the rear wingspan may be between about30 feet and about 40 feet.

The fixed wings 20, 24 may fold or pivot. For example, in one embodimenteach of the fixed wings 20, 24 has a chord 28, 30 and the fixed wingsare pivotally mounted on the fuselage 14 for selective movement betweena fixed-wing flight position, in which the respective chord extendsgenerally horizontally, and a rotary-wing flight position, in which therespective chord extends generally vertically. The forward flightposition of the fixed wings 20, 24 is shown by solid lines in FIG. 9 andgenerally indicated by reference arrow F and the vertical flightposition is shown by dashed lines and generally indicated by referencearrow V. The fixed wings 20, 24 may also be moved to intermediate flightpositions (not shown) between the forward and vertical flight positionswherein the respective wing chord 28, 30 is between horizontal andvertical.

As shown in FIG. 8, in this embodiment, the aircraft 10 further includesone or more power plants 32, 34 mounted on the airframe. The powerplants 32, 34 produce power in the form of hot high-pressure gas orexhaust during their operation. Although the power plants 32, 34 mayproduce other amounts of power without departing from the scope of thepresent disclosure, in some embodiments the power plants produce betweenabout 11,000 pounds and about 13,000 pounds of thrust. Although otherpower plants 32, 34 may be used without departing from the scope of thepresent disclosure, in one embodiment each power plant is a F404Turbofan available from General Electric Company of Cincinnati, Ohio.

The aircraft 10 also includes at least one rotor/wing, generallydesignated by 36, rotatably mounted on the aircraft by way of a driveshaft 38. The rotor/wing 36 includes a plurality of blades 40 extendingradially from a central hub 42 that is connected to the drive shaft 38to a blade tip 44. In one embodiment, the rotor/wing 36 has two primaryblades 40 extending from the hub 42 in opposite directions from eachother. Although the blades 40 may have other lengths between the hub 42and the respective blade tips 44, in one embodiment each blade has alength of between about 30 feet and about 35 feet. Because the blades 40and the drive shaft 38 do not need to be configured for routing exhaust,the blades and drive shaft can be thinner and lighter than the bladesand rotor mast of reaction-drive rotor/wing aircraft. The reduced weightand drag characteristics of the rotor/wing 36 improves aircraft 10performance and lowers power requirements compared to reaction-drivesystems. Although the blades 40 may have other maximum thicknesses 46without departing from the scope of the present disclosure, in oneembodiment each blade has a maximum thickness of between about 1 footand about 2 feet. Although the rotor blades 40 may be made of othermaterials, in one embodiment at least a portion of the blades are madeof a polymer composite.

The aircraft 10 has a rotation mode wherein the rotor/wing 36 is rotatedby the power plants 32, 34 and a fixed mode wherein the rotor/wing islocked to prevent rotor/wing rotation. In the rotation mode, therotor/wing 36 rotates to provide upward thrust to the aircraft 10. Theprimary fixed wings 20 are moved to their vertical flight position Vwhen the aircraft 10 is in the rotation mode so the primary fixed wingsminimally interfere with rotor 36 downwash and thus minimally inhibitthe production of upward thrust by the rotor. The rear fixed wings 24are also rotated to their vertical flight position when the aircraft 10is in the rotation mode so they minimally inhibit upward propulsion. Inthe fixed mode, the rotor/wing 36 is locked so the blades 40 extendlaterally to provide aerodynamic lift to the aircraft 10 during forwardflight. The aircraft 10 may also fly at intermediate flight modeswherein the aircraft is propelled at an angle between vertical andhorizontal. For example, an aircraft 10 transitioning between verticaland horizontal flight will fly at angles between vertical andhorizontal. The fixed wings 20, 24 are moved to their forward flightpositions F when the aircraft 10 is in the fixed mode and can assumeintermediate flight positions corresponding to intermediate flightmodes.

The aircraft 10 also includes a radial inflow turbine, generallydesignated by 48, mounted on the airframe 12 in fluid communication withthe power plants 32, 34 for receiving exhaust from the power plants. Theradial inflow turbine 48 is mechanically connected to the rotor/wing 36and converts exhaust from the power plants 32, 34 to mechanical powerfor rotating the rotor/wing during operation of the aircraft 10. Lossesincurred in converting the exhaust to mechanical power for rotating therotor/wing 36 are generally lower than the losses incurred between thepower plant(s) and the rotor/wing in a conventional reaction-driverotor/wing system. The higher efficiency of the radial inflow radialinflow turbine 48 system according to the present disclosure enableshigh performance and uses less power than is required for reaction-drivesystems. As shown in FIG. 10, the radial inflow radial inflow turbine 48includes a body or housing 50 forming a first inlet 52 and a secondinlet 54. As shown in FIG. 8, the first and second inlets 52, 54 are influid communication with the first and second power plants 32, 34,respectively. The turbine body 50 also forms a first aft outlet 56 and asecond aft outlet 58 downstream from the first and second inlets 52, 54,respectively.

In addition, the turbine body 50 forms an annular vortical plenum orchamber 60 in fluid communication with the inlets 52, 54 and outlets 56,58. As shown in FIG. 11, the vortical chamber 60 has an upper portion 62and a lower portion 64. Although the upper portion 62 of the vorticalchamber 60 may have other minimum radii 66 without departing from thescope of the present disclosure, in one embodiment the upper portion hasa minimum radius of between about 20 inches and about 22 inches.Although the lower portion 64 of the vortical chamber 60 may have othermaximum radii 68 without departing from the scope of the presentdisclosure, in one embodiment the lower portion has a maximum radius ofbetween about 7 inches and about 9 inches. The radial inflow radialinflow turbine 48 further includes a chamber outlet 70 downstream fromthe vortical chamber 60. Exhaust from the power plants 32, 34 passingthrough the vortical chamber 60 exits the radial inflow turbine 48 withreduced energy by way of the chamber outlet 70. Upon exiting the chamberoutlet 70, the exhaust flows into a low-energy conduit 72, as shown inFIG. 8.

The radial inflow turbine 48 also includes a hub 74 rotatably connectedto the turbine body 50 and a plurality of vanes 76 extending radiallyoutward from the hub. The hub 74 and the vanes 76 are positioned in theturbine vortical chamber 60. Each of the vanes 76 includes a top 78positioned in the upper portion 62 of the vortical chamber 60 and abottom 80 positioned in the lower portion 64 of the vortical chamber.Each vane 76 is pitched from its top 78 to its bottom 80. As will beappreciated by those skilled in the art, the pitch of the vanes 76creates an oblique surface 82 against which power plant 32, 34 exhaustis directed to cause the vanes 76 and hub 74 to rotate during operationof the aircraft 10 in the rotation mode. In one embodiment, each vane 76has a maximum radius 84 corresponding to the minimum radius 66 of theupper portion 62 of the vortical chamber 60 and a minimum radius 86corresponding to the maximum radius 68 of the lower portion 64 of thevortical chamber. The radial inflow turbine 48 further includes aturbine shaft 88 operatively connected to the turbine hub 74 and to therotor/wing drive shaft 38. In one embodiment, the rotor/wing drive shaft38 and the turbine shaft 88 are integrally formed. The turbine hub 74,the vanes 76, and the turbine shaft 88 rotate together and therotor/wing 36 is rotated by torque received from the turbine shaftduring operation of the aircraft 10.

As shown in FIG. 8, the aircraft 10 may include a gearbox 90 connectedto the turbine shaft 88 and the rotor/wing drive shaft 38 fortransmitting power transferred from the turbine shaft to the driveshaft. In one embodiment, the gearbox 90 is a reduction gearbox forreducing the power and rotational speed imparted to the drive shaft 38from the turbine shaft 88. In one embodiment, the gearbox 90 is aplanetary gearbox. Although other types of gearboxes 90 may be usedwithout departing from the scope of the present disclosure, in oneembodiment the gearbox is an accessory gearbox available from NorthstarAerospace Inc of Bedford Park, Ill. The gearbox 90 may have one or morestages and although the gearbox 90 may have other reduction ratioswithout departing from the scope of the present disclosure, in oneembodiment the gearbox has a reduction ratio of between about 7:1 andabout 9:1.

As shown in FIGS. 11 and 12, the radial inflow turbine 48 furtherincludes an inlet valve 92, 94 positioned within the turbine body 50adjacent to each inlet 52, 54. Although the inlet valves 92, 94 may beother types without departing from the scope of the present disclosure,in one embodiment each valve is a butterfly valve (also known as asliding door valve) or a ball valve. The inlet valves 92, 94 selectivelyallow power plant 32, 34 exhaust to pass through the radial inflowturbine 48 from the respective inlet 52, 54 to the corresponding aftoutlet 56, 58 for high-speed flight in the fixed mode or direct theexhaust through the vortical chamber 60 for flight in the rotation mode.For directing power plant 32, 34 exhaust through the vortical chamber60, the exhaust is first diverted from the respective inlet 52, 54generally upward into the upper portion 62 of the vortical chamber 60 bythe respective inlet valve 92, 94, then the exhaust flows generallyradially inward in the vortical chamber and generally downward throughthe vortical chamber and against the oblique surfaces 82 of the vanes76, as shown by arrow E in FIG. 12. As described above, the exhaustflowing against the oblique surfaces 82 of the vanes 76 causes the vanesand turbine hub 74 to rotate thereby rotating the turbine shaft 88, thedrive shaft 38, and the rotor/wing 36.

For embodiments having a single power plant (not shown), the radialinflow turbine 48 can be configured in a variety of ways. For example,the turbine 48 may include a sole inlet positioned at about a center ofan upstream end of the turbine for transferring exhaust from a singlepower plant to the vortical chamber and a sole outlet positioned atabout a center of a downstream end of the turbine. It is contemplatedthat in one embodiment (not shown), the exhaust from two or more powerplants are combined upstream from the turbine and enter the turbinethrough a sole turbine inlet.

As shown in FIGS. 8 and 9, the aircraft 10 further comprises a nozzle 96mounted on the airframe 12 adjacent to the aft end 18 of the fuselage14. The nozzle 96 is in fluid communication with the power plants 32, 34for receiving exhaust. The nozzle 96 may be operatively connected toeach aft outlet 56, 58 of the radial inflow turbine 48 for receivingpower plant 32, 34 exhaust exiting the aft outlets for high-speed flightin the fixed mode. For example, a high-energy conduit 98 (shown in FIG.8) may connect the aft outlets 56, 58 to the nozzle 96. The nozzle 96may also be operatively connected to the chamber outlet 70 for receivingexhaust during aircraft 10 operation. For example, the aircraft 10 mayfurther comprise a conduit valve 100 for selectively diverting exhaustflowing through the low-energy conduit 72 to the high-energy conduit 98and to the nozzle.

The nozzle 96 may be adjustable between multiple positions to providethrust in various directions. In one embodiment, the nozzle 96selectively directs exhaust to exit the aircraft 10 at a pre-selectedangle E (shown in FIG. 9) with respect to the airframe 12 within a rangeof angles extending from about horizontally rearward (i.e., θ is about0°), as shown by solid lines, and about vertically downward (i.e., θ isabout 90°), as shown by dashed lines. When the nozzle 96 is angledrearward, the exhaust exiting the aircraft 10 provides forward thrustand when the nozzle is angled downward, the exhaust provides upwardthrust. When the nozzle 96 is vectored to an angle θ between about 0°and 90°, the exiting exhaust provides thrust between forward and upwardaccording to the position of the nozzle. Aircraft 10 capable ofproviding vertical thrust from two locations of the aircraft arereferred to as two-poster aircraft. In the present disclosure,two-poster characteristics are present in the upward thrust provided atthe aft end 18 of the fuselage 14 by vectoring the nozzle 96 to an angleθ greater than zero and by the rotor/wing 36. Two-poster aircraft 10have enhanced flight performance abilities compared to single-posteraircraft because they can operate with a wider variety of centers ofgravity by controlling the amount of vertical thrust produced at eachposter. For embodiments of the aircraft 10 comprising multiple powerplants 32, 34 and corresponding turbine inlets 52, 54 and aft outlets56, 58, the aircraft may include a separate nozzle (not shown) in fluidcommunication with each aft outlet.

As shown in FIG. 8, the aircraft 10 also comprises a yaw control system,generally designated by 102, mounted on the airframe 12 adjacent to theaft end 18 of the fuselage 14. The yaw control system 102 is in fluidcommunication with the power plants 32, 34 for receiving exhaust foractively controlling yaw. Specifically, the yaw control system 102 isoperatively connected to the chamber outlet 70 by way of the low-energyconduit 72 for receiving power plant 32, 34 exhaust exiting the radialinflow turbine 48 through the chamber outlet during aircraft 10operation. The yaw control system 102 may also be operatively connectedto the aft outlets 56, 58 for receiving exhaust during aircraft 10operation. For example, the conduit valve 100 may be configured forselectively diverting exhaust flowing through the high-energy conduit 98to the low-energy conduit 72 during aircraft 10 operation. Yaw controlmay be needed to control aircraft 10 yaw during operation in therotation mode. Although the yaw control system 102 may be other typeswithout departing from the scope of the present disclosure, in oneembodiment (not shown) the yaw control system is a NOTAR® systemavailable from the Boeing Company of Chicago, Ill. NOTAR® is a federallyregistered trademark of the Boeing Company. In one embodiment, the yawcontrol system 102 includes right and left lateral outlets 104, 106connected by a valve 108. The yaw control system valve 108 is controlledto selectively direct exhaust received from the low-energy conduit 72 tothe right lateral outlet 104, to the left lateral outlet 106, or to bothlateral outlets to control 10 yaw during operation of the aircraft 10.

It will be appreciated that additional embodiments in accordance withthe present disclosure may be conceived, and that the invention is notlimited to the particular embodiments described above and shown in theaccompanying figures. For example, aspects of the various embodimentsshown in some of the figures may be selectively combined, included, orsubstituted for various aspects of other embodiments shown in otherfigures to create still other embodiments. For example, in someembodiments, the aircraft shown in FIG. 5 may be configured to include avectorable nozzle assembly (e.g. FIG. 4), or pivotable canards (e.g.FIG. 2-4), or a pivotable tail assembly (e.g. FIG. 2-4) to createadditional embodiments. Alternately, the vertical tail surfaces of thetail assemblies shown in FIGS. 1-7 may be eliminated and replaced with asingle vertical tail that extends generally upwardly from an aft portionof the fuselage to provide further embodiments. The embodiments shown inFIGS. 1-7 may be configured as single engine configurations ormultiple-engine configurations. Furthermore, one or more aspects of theembodiments shown in the accompanying figures may be modified or omittedcompletely to provide still other embodiments in accordance with thepresent disclosure.

When introducing elements of the present disclosure or the preferredembodiment(s) thereof, the articles “a”, “an”, “the”, and “said” areintended to mean that there are one or more of the elements. The terms“comprising”, “including”, and “having” are intended to be inclusive andmean that there may be additional elements other than the listedelements.

As various changes could be made in the above constructions withoutdeparting from the scope of the disclosure, it is intended that allmatter contained in the above description or shown in the accompanyingdrawings shall be interpreted as illustrative and not in a limitingsense. Thus, while preferred and alternate embodiments of the disclosurehave been illustrated and described, as noted above, many changes can bemade without departing from the spirit and scope of the invention.Accordingly, the scope of the invention is not limited by the disclosureof these preferred and alternate embodiments. Instead, the inventionshould be determined entirely by reference to the claims that follow.

What is claimed is:
 1. An aircraft comprising: a propulsion system; a rotor/wing assembly; and a radial inflow turbine assembly including a radial inflow turbine coupled to the rotor/wing assembly and configured to receive exhaust from the propulsion system during a rotational mode of operation and rotate the rotor/wing assembly, the turbine assembly configured to allow the exhaust to bypass the turbine and provide forward thrust during a second mode of operation.
 2. The aircraft according to claim 1 wherein the propulsion system includes at least one engine that produces an exhaust flow, and wherein the radial inflow turbine assembly is configured to extract work from the exhaust flow.
 3. The aircraft according to claim 1 wherein the propulsion system includes a vectorable nozzle assembly configured to controllably direct an exhaust flow between an approximately aftward direction and an approximately downward direction.
 4. The aircraft according to claim 1, further comprising a yaw control system configured to provide a lateral force by exhausting a portion of an exhaust flow in a lateral direction.
 5. The aircraft according to claim 1 further comprising a fuselage; wherein a rotational axis of the rotor/wing assembly extends generally upwardly with respect to the fuselage, and wherein the radial inflow turbine assembly is configured such that a drive portion of an exhaust flow is received into a body along an approximately radial direction with respect to the rotational axis, is deflected by a drive portion, and exits from the body approximately along the rotational axis.
 6. The aircraft according to claim 5 wherein the drive assembly further includes a drive shaft projecting upwardly with respect to the fuselage and connected to the rotor/wing assembly; and wherein the radial inflow turbine assembly includes a turbine blade assembly coupled to the drive shaft and including a plurality of vanes that provide a rotational force when subject to the exhaust flow.
 7. The aircraft according to claim 1 further comprising a fuselage and canards extending outward laterally from the fuselage; wherein each canard is pivotable between a first position configured to provide a lift force in a fixed-wing mode of operation, and a second position configured to minimize a downward force in a rotary-wing mode of operation.
 8. The aircraft according to claim 1 further comprising a tail assembly pivotable between a first tail position configured to provide a lift force in the second mode of operation, and a second tail position configured to minimize a downward force in the first mode of operation.
 9. A drive assembly for an aircraft having a propulsion system and a rotor/wing assembly, the drive assembly comprising: a radial inflow turbine assembly operatively coupled to the rotor/wing assembly and to the propulsion system and configured to extract work from the propulsion system to selectively rotate the rotor/wing assembly.
 10. The drive assembly according to claim 9 wherein the propulsion system includes at least one engine that produces an exhaust flow, and wherein the radial inflow turbine assembly is configured to extract work from the exhaust flow.
 11. The drive assembly according to claim 9 wherein the propulsion system includes at least one engine that produces an exhaust flow, and wherein the rotor/wing assembly rotates about a rotational axis, and wherein the radial inflow turbine assembly is configured such that a drive portion of the exhaust flow is received into a body along an approximately radial direction with respect to the rotational axis, is deflected by a drive portion, and exits from the body approximately along the rotational axis.
 12. The drive assembly according to claim 11 wherein the radial inflow turbine assembly is configured to be coupled to the rotor/wing assembly by a drive shaft, and wherein the radial inflow turbine assembly includes a turbine blade assembly configured to be coupled to the drive shaft and including a plurality of vanes that provide a rotational force when subject to the exhaust flow.
 13. The drive assembly according to claim 9, further comprising means for allowing the exhaust gas to bypass the turbine and provide forward thrust during a second mode of operation.
 14. The aircraft according to claim 1, wherein the turbine assembly further includes a turbine body that forms a vortical chamber; wherein the radial inflow turbine is located in the vortical chamber; and wherein the turbine assembly further includes at least one inlet valve positioned within the body to selectively allow the exhaust to bypass the radial inflow turbine during the first mode and to direct the exhaust into the vortical chamber during the second mode.
 15. The aircraft according to claim 1, wherein the turbine is coaxial with a hub of the rotor/wing assembly.
 16. The aircraft according to claim 1, wherein the propulsion system includes first and second engines arranged side-by-side, and wherein the turbine assembly includes a housing for the turbine, the turbine housing having first and second inlets in fluid communication with the engines.
 17. The aircraft according to claim 1, wherein the radial inflow turbine includes top vane portion and lower vane portion, the top vane portion having greater diameter than lower vane portion; and wherein the turbine assembly further includes a housing that forms a vortical chamber, the housing having an upper housing portion for housing the top vane portion and a lower housing portion for housing the lower vane portion.
 18. The aircraft according to claim 17, wherein the housing is configured to direct exhaust gas from the propulsion system into the upper housing portion and then into the lower housing portion. 